Radiator using thermal control coating

ABSTRACT

A spacecraft has internal structure which generates heat, and a radiator element in thermal communication with the spacecraft internal structure. The radiator element has a radiating surface and a coating on the radiating surface including a white thermal control paint. The paint has an initial solar absorptance of not greater than 0.16 and an initial infrared emittance of not less than about 0.80. The standard end-of-life projected solar absorptance is not more than about 0.33 for a paint with an inorganic binder, and the standard end-of-life projected solar absorptance is not more than about 0.6 for a paint with an organic binder.

BACKGROUND OF THE INVENTION

This application is a continuation in part of allowed application Ser.No. 08/819,747, filed Mar. 18, 1997, which became U.S. Pat. No.5,884,868 issued on Mar. 23, 1999.

This invention relates to radiators, and, more particularly, toradiators used in spacecraft.

A radiator is designed to radiate heat to its surroundings. Spacecraftwhose occupants, electronics, or power sources generate large amounts ofheat employ one or more radiators to transfer the generated heat fromthe interior of the spacecraft to free space and to reflect incidentheat from solar radiation exposure. The radiators also aid indissipating static electricity on the surface of the spacecraft. Theradiators are necessary to prevent heating of the interior of thespacecraft to unacceptably high levels. For some spacecraft such aslarge communications satellites that generate and utilize large amountsof power, removal of excess heat is a significant factor in the designof the spacecraft, and large amounts of radiator surface are required.

In a commonly used construction of a spacecraft radiator, the radiatingsurface is formed of a large number of individual mirror-like radiators.A single larger mirror-like radiator is not used because of thelikelihood that it would crack due to thermal strains. Each mirror-likeradiator is 1-2 inches on a side. Each mirror-like radiator is formed ofa ceramic-glass substrate about 0.002-0.010 inches thick that is coatedon the inwardly facing surface with a metallic silver coating. Themetal-coated mirror has a relatively low solar absorptance and arelatively high infrared emittance, so that heat is effectively radiatedaway without absorbing excessive energy from incident sunlight. Theopposite, outwardly facing surface of the ceramic-glass substrate iscoated with a layer of transparent indium-tin-oxide that serves todissipate static charges. The metallic silver-coated inwardly facingsurface is bonded to the underlying structure with a silicone adhesive.

This radiator construction is operable and widely used on communicationssatellites. However, the fabrication of the spacecraft using theindividual mirror-like radiators is a time-consuming, expensive process.Hundreds or thousands of individual mirror-like elements are fabricatedby deposition processes and then individually attached to the underlyingsupport surface. Because the glass ceramic mirror substrates are thinand large in lateral extent relative to their thickness, they arefragile and easily broken during fabrication, assembly, or service.

There is accordingly a need for an improved approach to the constructionof radiators used in spacecraft and for other applications. The presentinvention fulfills this need, and further provides related advantages.

SUMMARY OF THE INVENTION

The present invention provides a radiator and a spacecraft whichutilizes such a radiator. The radiator has excellent performance in thespace environment. It is fabricated and installed in the spacecraft muchless expensively and more quickly than a radiator using conventionalmirror-like radiating elements. The weight of the radiator is reduced ascompared with the mirror-like radiator, an important consideration forlaunching the spacecraft from earth. The radiator of the invention ismore robust than a mirror-like radiator.

In accordance with the invention, a radiator comprises a structure whichgenerates heat, and a radiator element in thermal communication with thestructure. The radiator element comprises a radiating surface, with acoating on the radiating surface comprising a white thermal controlpaint. The paint has an initial solar absorptance of not greater than0.16, preferably not greater than about 0.14 and more preferably notgreater than about 0.10, and an initial infrared emittance of not lessthan about 0.80. (The “initial” properties are those measured after thepaint is applied and dried, but before any substantial exposure to aspace environment.) The structure which generates heat is preferably theinternal structure of a spacecraft such as a satellite.

The paint used in the coating on the radiating surface of the radiatoris formed of white pigment particles in a binder. The preferred pigmentparticles have a composition of Zn[xAl(1−x)Ga]₂O₄(δIn), termed a zincaluminate gallate, where the value of x is from 0 to 1 and the value ofδ is from 0 to about 0.2. The pigment particles may also be azinc-containing pigment such as the doped or undoped ZnO pigmentdisclosed in U.S. Pat. No. 5,094,693, whose disclosure is incorporatedby reference. The binder is preferably an inorganic binder such as asilicate, and preferably potassium silicate, but an organic binder mayalso be used for less-demanding applications. The weight ratio ofpigment to binder is preferably from about 3:1 to about 5:1, but may beless than about 3:1. The paint thickness is preferably from about 0.003to about 0.006 inches, after drying. The zinc aluminate gallate paintwith a potassium silicate binder has an initial solar absorptance α ofless than 0.10, and aluminum-doped zinc oxide paint with a potassiumsilicate binder has an initial solar absorptance of about 0.13-0.18.

The paint is applied to the radiating surface of the radiator usingconventional painting techniques such as brushing or spraying, or bynon-vehicle painting techniques such as plasma spray.

Many white coatings have been used for the exterior portions ofspacecraft other than the radiators. During extended exposure to thespace environment of ultraviolet radiation, gamma radiation, electrons,and protons, the initially white coatings become yellow and then gray astheir solar absorptance rises. As the solar absorptance rises, thecoatings become less efficient thermal coatings, because they absorbincreasing amounts of solar energy. This loss of efficiency typicallytakes at least several years and is not a concern for short-livedspacecraft or satellites.

The gradually increasing solar absorptance with increasing exposure isof relatively little significance for the non-radiator portions of theexterior of the spacecraft, even those spacecraft such as communicationssatellites that spend many years in space. These non-radiator regionsare not designed to dissipate larger amounts of heat than they absorb,and in many cases do not face the sun. If the solar absorptance forthese portions of the spacecraft rises, there is relatively littleeffect on the total heat balance of the spacecraft. The radiators, onthe other hand, must dissipate large amounts of interiorly generatedheat. Due to the required orientation of the antennas and solar cells ofthe spacecraft, the radiators often must face toward the sun. If thesolar absorptance of the radiators increases significantly over theoperating life of the spacecraft, the radiators begin to absorb largeamounts of heat and consequently radiate the interiorly generated heatless efficiently. The thermal balance of the spacecraft is degraded. Thetemperature of the interior of the spacecraft begins to rise, andeventually exceeds the acceptable operating temperatures.

Thus, the radiators of the spacecraft have different operatingconditions and requirements than the exterior non-radiator walls of thespacecraft. The coatings that have long been used on the non-radiatorexterior walls of the spacecraft could not be used for the radiatorsbecause of their inadequate long-term stability in the spaceenvironment. The development of the white coatings with low initialvalues of solar absorptance and high initial values of solar emittancehas made possible the present approach to an improved radiator, becausethe initial values of solar absorptance are sufficiently far removedfrom the maximum permissible value after extended exposure that thegradual increase in solar absorptance does not result in theinoperability of the radiator.

At the present time, there is no known approach for preventing thegradual increase in the solar absorptance that occurs over time.Instead, the initial solar absorptance must be selected to besufficiently low such that, after the gradual increase in solarabsorptance over time and exposure, the paint still has an acceptablesolar absorptance. In the present invention, the standard end-of-lifesolar absorptance is not greater than about 0.33 when an inorganicbinder is used, most preferably not greater than about 0.30. In order toachieve these standard end-of-life solar absorptances, the initial solarabsorptance is not greater than 0.16, preferably not greater than about0.14, and most preferably not greater than about 0.10.

The radiator of the invention provides an important advance inspacecraft systems and particularly in heat dissipation and reflectionby the spacecraft radiator. The radiator is effective in dissipatingheat to the surroundings and in reflecting incident solar radiation,both initially and after extended exposure to the space environment. Itis much more readily fabricated and installed than a conventionalradiator having a mirror-like structure. The fabrication andinstallation are also accomplished much faster than in the priorapproach. The latter is an important consideration in the highlycompetitive business of communications satellites, where delivery timeson the order of 12 months are becoming common, as compared with normaldelivery times on the order of 24 months or longer in the past. Otherfeatures and advantages of the present invention will be apparent fromthe following more detailed description of the preferred embodiment,taken in conjunction with the accompanying drawings, which illustrate,by way of example, the principles of the invention. The scope of theinvention is not, however, limited to this preferred embodiment.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic perspective view of a communications satellite;

FIG. 2A is an enlarged schematic sectional view of a first embodiment ofa portion of the satellite of FIG. 1, taken along line 2—2;

FIG. 2B is an enlarged schematic sectional view of a second embodimentof a portion of the satellite of FIG. 1, taken along line 2—2;

FIG. 3 is an enlarged view of the coating applied to the radiatingsurface of the radiator; and

FIG. 4 is a block flow diagram of a method for practicing the invention.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 is a schematic view of a spacecraft, here a communicationssatellite 20, having radiators 22 of three types affixed to the body ofthe satellite. A first radiator 22 a overlies a portion of an externalwall 24 of the satellite and is disposed in fixed relation to theremainder of the satellite. A second radiator 22 b is supported in fixedrelation to a projection 26 that contains the batteries of thesatellite. A third radiator 22 c is mounted in a deployable fashion fromthe satellite, on a hinge mechanism 28. Any or all of these radiators 22a, 22 b, and 22 c may be made using the approach of the invention. Ingeneral, the radiators 22 are distinct from other exteriorly facingportions of the spacecraft such as the skin structure, the solar cells,and the antennas, in that the radiators 22 are made of materials withhigh thermal radiation properties and are in thermal communication withinternal heat sources of the spacecraft by direct contact or byhigh-thermal-conductivity thermal transfer media, and are intentionallydesigned to radiate large amounts of heat. The structure of thespacecraft radiator will be discussed in more detail in reference toFIG. 2A.

FIGS. 2A and 2B are schematic sectional views of two designs of theradiator 22. The radiator 22 is positioned adjacent to the external wall24 (sometimes termed the “skin”) of the satellite 22, and in someinstances may be flush with the external wall 24. The radiator 22includes a radiator body 30 having an outwardly facing radiating surface32 and an inwardly facing surface 34. The radiator body 30 is made of amaterial that is a good thermal conductor, such as aluminum or aluminumalloy, or a relatively high-thermal-conductivity composite material suchas graphite/epoxy composite material. In one approach, the radiator bodyis a honeycomb structure with an outwardly facing face sheet bonded tothe honeycomb. The external wall 24, on the other hand, is typically nota good thermal conductor and is insulated from internal heat sources.

The inwardly facing surface 34 is in thermal communication with a heatsource 36 that is part of the internal structure of the spacecraft. Theheat source 36 may be any heat source such as, for example, a battery,an electronics unit, or a crew compartment. The inwardly facing surface34 may be in thermal communication with the heat source 36 through athermal transfer medium 38, as shown in FIG. 2A. The thermal transfermedium has a first end 40 in thermal communication with the heat sourceand a second end 42 in thermal communication with the inwardly facingsurface 34 of the radiator body 30. The thermal transfer medium 38 maybe any thermal conductor, such as a metallic or nonmetallic strip, aheat pipe, or other medium. Equivalently for the present purposes and asillustrated in FIG. 2B, the heat source 36 may be in direct thermalcontact with the inwardly facing surface 34 of the radiator body 30.

The outwardly facing radiating surface 32 of the radiator body 30 has acoating 44 applied thereto. The coating 44 includes a white thermalcontrol paint having a structure generally illustrated in FIG. 3. Thecoating 44 may cover all or only a portion of the surface 32, althoughin practice it would normally cover the entire surface 32 as shown inFIGS. 2A and 2B.

The coating 44 is a paint comprising a plurality of pigment particles 46mixed with a binder 48. The pigment particles are white in color. Theparticles are preferably of a composition Zn[xAl(1−x)Ga]₂O₄(δIn), wherethe value of x is from 0 to 1 and the value of δ is from 0 to about 0.2,which has a spinel crystal structure. The particles may instead be thedoped or undoped zinc oxide particles disclosed in U.S. Pat. No.5,094,693, which has a wurtzite crystal structure, or other operableparticles.

The binder is an inorganic or organic material. The preferred inorganicbinder is a silicate, most preferably potassium silicate, but othersilicates such as sodium silicate may also be used. The binder mayinstead be an organic material such as dimethyl silicone,poly(dimethyl-siloxane), polyurethane, polyamide, or polyurea. Theinorganic binder is preferred for use in radiator paints having the mostdemanding exposure conditions, such as geosynchronous communicationssatellites that remain in orbit for many years. For these applications,spacecraft designs require that the solar absorptance of the radiatorpaint not exceed about 0.33, preferably about 0.30, after extendedexposure to the spacecraft environment. The organic binder may be usedin less-demanding applications such as medium earth orbit communicationssatellites that typically orbit in the 6,000-10,000 mile altitude range.For these less-demanding applications, spacecraft designs require thatthe solar absorptance of the radiator paint not exceed about 0.6 afterextended exposure.

For both the inorganic and the organic binders, the weight ratio ofpigment to binder is preferably from about 3:1 to about 5:1, althoughthe coating is operable for values outside this range. The thickness ofthe coating 44 is preferably from about 0.003 to about 0.006 inchesafter drying to a solid form, although the coating is operable forvalues outside this range. In this range, the thicker the coating of athermal control paint, the lower is the solar absorptance but thegreater the weight of the coating. However, increasing thicknesses aboveabout 0.006 inches increase the weight without corresponding reductionsin solar absorptance.

FIG. 4 depicts a preferred method for preparing the particles 46, forpreparing the paint material used in the coating layer 44, and forpainting the radiating surface 32 of the radiator body 30.

If commercially available, the pigment particles may be purchased. Ifnot, they may be prepared by the following procedure. The components ofthe particles are provided and mixed together, numeral 70. In thepreferred formulation procedure, the readily available components ZnO,Al₂O₃, Ga₂O₃, and In₂O₃ are used as starting materials. Thus, to prepareZnAl₂O₄, the appropriate amounts of ZnO and Al₂O₃ are mixed together. Toprepare ZnGa₂O₄, the appropriate amounts of ZnO and Ga₂O₃ are mixedtogether. To prepare Zn[xAl (1−x)Ga]₂O₄, the appropriate amounts of ZnO,Al₂O₃, and Ga₂O₃ are mixed together. If any of these compositions is tobe doped with indium, as is normally the case, the appropriate amount ofIn₂O₃ is added to the mixture. A mixing medium, which later is removed,may be added to promote the mixing of the components. Preferably, wateris used as the mixing medium.

The components and the mixing medium are milled together to form amechanical mixture, numeral 72. After milling is complete, the mixingmedium is removed by evaporation, numeral 74. The dried mixture is firedto chemically react the components together, numeral 76, at atemperature that is preferably in the range of from about 1000° C. toabout 1300° C. A preferred firing treatment is 1160° C. for 6 hours, inair. After cooling, the agglomerated mass resulting from the firing islightly pulverized, as with a mortar and pestle, numeral 78. Theresulting particulate has a size range of from about 0.1 micrometer toabout 5 micrometers. The preparation of the particulate pigment iscomplete.

The paint is prepared by providing the particulate material, prepared asdescribed above or otherwise. The binder is provided, numeral 80, toadhere the particles together in the final product. The binder isselected to provide good adherence of the particles to each other and ofthe particles to the underlying substrate, with acceptable physicalproperties. For example, the binder must withstand the environment towhich the paint is exposed, such as space.

The preferred inorganic binder for demanding applications is potassiumsilicate. The binder is present in an operable amount. In a typicalcase, the binder is present in an amount such that the ratio, by weight,of the pigment to the binder is from about 3:1 to about 5:1. If theratio is less than about 3:1, the paint is operable but tends to betranslucent after drying, so that the initial value of the solarabsorptance is too high. In that case, a second coat or a thicker firstcoat of the paint may be required. If the ratio is more than about 5:1,the critical pigment volume concentration (CPVC) is exceeded, the painthas insufficient mechanical strength, and the paint falls apart whendried.

The pigment particles and the binder are jointly selected so that thefinal dried paint has an initial solar absorptance of not greater than0.16, preferably not greater than about 0.14, and most preferably notgreater than about 0.10, and an initial infrared emittance of not lessthan about 0.80. A high infrared emittance of not less than about 0.8 isrequired in order to radiate heat away from the radiator. If the initialinfrared emittance is less than about 0.8, the radiator will noteffectively radiate heat away. The infrared emittance does not degradesubstantially with extended exposure to the space environment. If theinitial solar absorptance of the paint is in excess of 0.16, it will bedifficult or impossible to maintain the standard end-of-life projectedsolar absorptance within the required limits. Preferably, the initialsolar absorptance of the paint is less than about 0.14 and mostpreferably less than about 0.10, so as to ensure that the standardend-of-life projected solar absorptance is less than the upper limitsdictated by spacecraft design.

A low solar absorptance is required so that the radiator is notexcessively heated by incident thermal energy when the radiator isfacing the sun. White paints tend to age and become yellow and then graywith extended periods of exposure to a space environment. As the paintbecomes gray, its solar absorptance increases so that the efficiency ofthe radiator falls. It is therefore preferred that the paint with aninorganic binder have a standard end-of-life projected solarabsorptance, as determined by standard accelerated exposure testing, ofno greater than about 0.33, preferably about 0.30. It is preferred thatthe paint with an organic binder have a standard end-of-life projectedsolar absorptance, as determined by standard accelerated exposuretesting, of no greater than 0.60. If these values are substantiallyexceeded, the radiator does not function sufficiently well toward theend of its life that heat is efficiently removed from the interior ofthe spacecraft, and the interior components overheat. (The “standardend-of-life projected solar absorptance” is that measured after thepaint has been applied and dried, and thereafter subjected to a standardaccelerated exposure test of the type discussed next.)

It is not practical to test the paint under long-term space conditionsbecause many years of exposure is required in order to assess theend-of-life properties. Instead, well-known accelerated tests are usedto simulate the effect on the coating of long-term space exposure. Inthe accelerated test used to determine the standard end-of-lifeprojected solar absorptance to determine whether a paint falls withinthe scope of the present invention, the paint is tested by simultaneousor serial exposure to ultraviolet radiation, electrons, and optionallyprotons according to the following test protocols or their effectiveequivalents. The test protocols are as follows for both the presentpaint having the inorganic binder and the paint having the organicbinder. Ultraviolet exposure testing is accomplished by exposure for1000 hours to ultraviolet light having 1-3 times, preferably 1-2 times,and most preferably 1 time the ultraviolet intensity of the sun.Electron exposure testing is accomplished with a first exposure of1×10¹⁸ electrons per square centimeter at an energy of 35 keV (thousandsof electron volts) and a flux of 6×10⁹ electrons per squarecentimeter-second, and a second exposure of 1×10¹⁸ electrons per squarecentimeter at an energy of 100 keV and a flux of 6×10⁹ electrons persquare centimeter-second Proton exposure testing is accomplished with afirst exposure of 3×10¹⁶ protons per square centimeter at an energy of45 keV and a flux of 5×10¹³ protons per square centimeter-second, asecond exposure of 1×10¹⁶ protons per square centimeter at an energy of90 keV and a flux of 1×10¹³ protons per square centimeter-second, athird exposure of 5×10¹⁵ protons per square centimeter at an energy of160 keV and a flux of 5×10¹² protons per square centimeter-second, and afourth exposure of 3×10¹⁵ protons per square centimeter at an energy of300 keV and a flux of 3×10¹² protons per square centimeter-second.

The solar absorptances measured both before and after this standardend-of-life exposure are determined, and the solar absorptance afterexposure is taken as the measure of the standard end-of-life projectedsolar absorptance (corresponding to about 15 years of exposure ingeosynchronous orbit). For example, in the case of the Zn[xAl(1−x)Ga]₂O₄particles and preferred potassium silicate binder, the initial solarabsorptance of a paint having an inorganic binder is less than 0.10, andtypically about 0.06. After exposure, the solar absorptance is higher,but not greater than about 0.33, and preferably not greater than about0.3. In the case of the Zn[xAl (1−x)Ga]₂O₄ particles in an organicbinder, the initial solar absorptance is about 0.15. After exposure, thesolar absorptance is higher, but less than about 0.6. From this andother testing, the maximum acceptable initial solar absorptance isestimated to be 0.16 in order to achieve the desired standardend-of-life projected solar absorptance. If the initial solarabsorptance is greater than 0.16, the standard end-of-life projectedsolar absorptance will be unacceptably great. An initial solarabsorptance no greater than about 0.14 is preferred, to provide a marginof error in achieving the desired standard projected end-of-life solarabsorptance. An initial solar absorptance of no greater than about 0.10is most preferred, as it produces a lower standard end-of-life projectedsolar absorptance. Initial solar absorptances in the range of about 0.10to 0.16 are acceptable. The standard end of life absorptancerequirements, as discussed above (i.e., 0.33 maximum, preferably nogreater than 0.30) should also be met.

The mixture of pigment and binder is ordinarily a solid, and a paintvehicle may be added to form a solution or a slurry that may be appliedusing conventional painting techniques, numeral 82. One preferred paintvehicle is water, which does not have adverse environmental impacts whenlater evaporated. Organic paint vehicles such as xylene and naphtha mayalso be used. The amount of the paint vehicle is selected to provide aconsistency that will permit application of the paint by the desiredapproach. For example, application by spraying requires the use of moreof the paint vehicle than application by brush or roller.

The paint may instead be applied by a technique where no vehicle isused, and in that case the step 82 is omitted.

The particles, binder, and paint vehicle (where present) are mixedtogether and milled together, numeral 84, to form a liquid paintformulation in which the particles do not rapidly separate. There may besome separation over extended periods of time, but the paint is normallystirred or agitated just before or at the time of application. Thepreparation of the paint is complete.

The paint is applied by providing the radiator body 30 to be coated asthe substrate, numeral 86, and cleaning the substrate, numeral 88. Thereis no known limitation on the type of substrate, but for a radiator thesubstrate is a good thermal conductor such as a metal. A preferred metalis aluminum or an aluminum alloy. The surface of the substrate iscleaned by any operable technique, such as washing and scouring in adetergent solution, rinsing in tap water, rinsing in de-ionized water,and drying in air. A composite material having sufficiently good thermalconductivity such as a graphite/organic resin (e.g., epoxy) compositeradiator substrate may instead be used.

The paint is applied to the surface of the substrate, numeral 90. At theoutset of the application, the surface of the substrate may be primed toimprove the adhesion of the paint. Priming is preferred for applicationof the paint containing an inorganic binder to metallic surfaces such asaluminum. Preferably, the priming, if used, is accomplished by rubbing asmall amount of the paint into the surface using a clean cloth, toachieve good contact to the surface.

The paint layer is thereafter applied by any operable technique, withspraying being preferred. As indicated earlier, the amount of paintvehicle present in the paint is selected to permit application by thepreferred approach. At this point, the paint is a thin film of a liquidmaterial. The paint may also be applied by a plasma spray technique orthe like wherein the mixture of pigment and binder is supplied to aheated region such as a plasma and directed toward the substrate. Theplasma-heated mixture of pigment and binder strikes the substrate andsolidifies thereon.

The paint is dried as necessary to leave a thin film of a solidmaterial, numeral 92. For a paint with an organic binder, the drying ispreferably accomplished at ambient temperature with a 35 percent orgreater humidity and for a time of 7 days. For a paint with an inorganicbinder, any humidity level is acceptable. Drying removes the paintvehicle by evaporation. Additionally, the drying step may accomplish adegree of curing of any curable components, as where a curable organicor inorganic binder is used. The paint layer is preferably from about0.003 to about 0.006 inches thick after drying. The painting iscomplete.

If the radiator 22 is manufactured as a separate element, it isthereafter affixed to the wall 24 and placed into thermal contact withthe satellite internal structure that generates heat, as shown in FIGS.2A and 2B.

Although a particular embodiment of the invention has been described indetail for purposes of illustration, various modifications andenhancements may be made without departing from the spirit and scope ofthe invention. Accordingly, the invention is not to be limited except asby the appended claims.

What is claimed is:
 1. A spacecraft, comprising: a spacecraft satelliteinternal structure which generates heat; and a radiator element inthermal communication with the spacecraft internal structure whichgenerates heat, the radiator element comprising a radiating surface, anda coating on the radiating surface comprising a white thermal controlpaint having a thickness of from about 0.003 to about 0.006 inches, thepaint comprising a mixture of zinc-containing pigment particles and abinder and having an initial solar absorptance of not greater than 0.16and an initial infrared emittance of not less than about 0.80.
 2. Aspacecraft, comprising: a spacecraft internal structure which generatesheat; and a radiator element in thermal communication with thespacecraft internal structure which generates heat, the radiator elementcomprising a radiating surface, and a coating on the radiating surfacecomprising a white thermal control paint having an inorganic binder, thepaint having a standard end-of-life projected solar absorptance of notmore than about 0.33.
 3. The spacecraft of claim 2, wherein the binderis potassium silicate.
 4. A spacecraft, comprising: a spacecraftinternal structure which generates heat; and a radiator element inthermal communication with the spacecraft internal structure whichgenerates heat, the radiator element comprising a radiating surface, anda coating on the radiating surface comprising a white thermal controlpaint having an organic binder, the paint having a standard end-of-lifeprojected solar absorptance of not more than about 0.6.
 5. Thespacecraft of clam 4, wherein the binder is selected from the groupconsisting of dimethyl silicone, poly(dimethyl-siloxane), polyurethane,polyamide, and polyurea.
 6. A spacecraft, comprising: a spacecraftinternal structure which generates heat; and a radiator element inthermal communication with the spacecraft internal structure whichgenerates heat, the radiator element comprising a radiating surface, anda coating on the radiating surface comprising a white thermal controlpaint, the paint having an initial solar absorptance of not greater than0.16 and an initial infrared emittance of not less than about 0.80. 7.The spacecraft of claim 6, wherein the spacecraft is a satellite.
 8. Thespacecraft of claim 6, wherein the radiator element is disposed in fixedrelation to the spacecraft internal structure.
 9. The spacecraft ofclaim 6, wherein the radiator element is deployable between a firstposition in relation to the spacecraft internal structure and a secondposition in relation to the spacecraft internal structure.
 10. Thespacecraft of claim 6, wherein the radiating surface comprises amaterial selected from the group consisting of aluminum, an aluminumalloy, and a composite material.
 11. The spacecraft of claim 6, whereinthe paint has an initial solar absorptance of from about 0.10 to 0.16.12. The spacecraft of claim 6, wherein the paint has an initial solarabsorptance of not greater than about 0.14.
 13. The spacecraft of claim6, wherein the coating has a thickness of from about 0.003 inches toabout 0.006 inches.
 14. The spacecraft of claim 6, wherein the paintcomprises zinc-containing pigment particles.
 15. The spacecraft of claim6, wherein the spacecraft further includes a thermal transfer mediumhaving a first end in thermal communication with the spacecraft internalstructure which generates heat and a second end in thermal communicationwith the radiator element.
 16. The spacecraft of claim 6, wherein thespacecraft thermal structure which generates heat is in direct physicalcontact with the radiator.
 17. The spacecraft of claim 6, wherein thepaint comprises an organic binder.
 18. The spacecraft of claim 17,wherein the binder is selected from the group consisting of dimethylsilicone, poly(dimethyl-siloxane), polyurethane, polyamide, andpolyurea.
 19. The spacecraft of claim 17, wherein the paint has astandard end-of-life projected solar absorptance of not more than about0.6.
 20. The spacecraft of claim 6, wherein the paint comprises aninorganic binder.
 21. The spacecraft of claim 20, wherein the binder ispotassium silicate.
 22. The spacecraft of claim 20, wherein the painthas a standard end-of-life projected solar absorptance of not more thanabout 0.33.